Secondary fluid injection thrust vectoring methods and apparatus



D c- 7, 1955 1.. T. BANKSTON SECONDARY FLUID INJECTION THRUST VECTORING METHODS AND APPARATUS Filed Aug. 6, 1962 INVENT 'OR. LESTER T. BANKSTON LIQUID GUIDANCE CONTROL %/7% ATTORNEY.

United States Patent 3,221,498 SECONDARY FLUID INJECTION THRUST VEC- TORING METHODS AND APPARATUS Lester T. Banlrston, Oxnard, Calif., assiguor to the United States of America as represented by the Secretary of the Navy Fiied Aug. 6, 1962, Ser. No. 215,252 6 Claims. (Cl. 60-3554) (Granted under Title 35, U.S. Code (1952), see. 266) The invention described herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.

This invention relates to jet propulsion apparatus and more particularly to improvements for controlling the direction of thrust produced by a nozzle to thereby control the direction of movement of a missile or the like which is propelled by a jet motor.

Since the advent of the jet motor propelled vehicle which operates in the environment beyond the atmosphere, various directional control systems have been devised which in various ways cause controlled deflection of the axis of thrust away from the direction of vehicle movement to thereby produce a lateral force which moves the vehicle in the desired direction. Devices of this type include movable jet vanes disposed in the thrust nozzle, universally gimballed nozzles, gas deflectors disposed rearwardly of the nozzle exit, sometimes referred to as jetavators, and various other devices which embody movable parts disposed in the hot gas stream and which require complicated and unreliable mechanical or other actuating devices. As this art developed, an entirely new concept evolved for achieving directional control by the discovery of the effects of injection of air or other gas into an exhaust nozzle, as exemplified by U.S. patent to Wetherbee, 2,943,821. Subsequent to the foregoing development it was discovered by Walker, U.S. Patent 2,916,873, that liquid could be vaporized by the exiting hot gases and produce a similar result. The present invention, as will subsequently more fully appear, involves injection of liquid into an exhaust gas stream but, unlike the Walker concept, does not depend for its operation upon vaporization, the liquid remaining in a liquid state in the exhaust stream. The present invention is acknowledged as being prior art in patent application of Richard D. Mc- Kee, Serial No. 155,217, filed November 27, 1961, which application relates to certain improvements over the present invention, particularly in the specific valving.

The principal objects of the invention are to provide method and apparatus for controlling direction of thrust of a fixed nozzle by controlled injection of liquid into the exhaust stream.

Further objects, advantages and salient features will become more apparent from a consideration of the description to follow, the appended claims, and the accompanying drawing, in which:

FIG. 1 is a side elevation of a rocket propelled missile embodying the present invention;

FIG. 2 is an enlarged section taken on line 22, FIG.

FIG. 3 is a section taken on line 33, FIG. 2, and diagrammatically illustrating certain parts not shown in in FIG. 2;

FIG, 4 is a section taken on line 4-4, FIG. 5, a modification for effecting roll control; and

FIG. 5 is a section taken on line 5-5, FIG. 4.

Referring now to the drawing, the subject of the invention comprises a conventional nozzle 10, such as of the converging-diverging de Laval type having a longitudinal axis 11 aligned with longitudinal axis 12 of a jet propelled 3,221,498 Patented Dec. 7, 1965 missile 13, and which is provided with a plurality of angularly spaced discharge orifices 14 in its divergent cone through which liquid may be selectively injected into the nozzle. Suitable conduits 16, provided with selectively operable valves 17 control the rate of liquid injection through the various discharge orifices. Any suitable pressurized liquid supply reservoir 18 may be employed for delivering the liquid to the various valves. As will be understood, the valves may be opened and closed, or partially opened for a rate and duration of discharge in accordance with the demands of any guidance control system 19 which senses deviation of the missile from its desired course and effects operation of the valves to provide necessary correction back to the desired course, such guidance systems being conventional and. well known.

It has been found that the side force produced is the result of two effects. First, the fluid bump or wedge in the exhaust stream produces an oblique shock front 20 through which the stream of gases is turned. Second, the injected liquid produces an area at the exit plane of the nozzle which has higher unit momentum than produced without injection. Both of these effects contribute to the side force. It has also been found that side force is increased if injection is effected at a point along the nozzle with higher mach numbers, that is, nearer the exit of the nozzle. This is limited, however, since if injection is too near the nozzle exit the oblique shock will extend rearwardly beyond the nozzle and become ineffective. It would appear, therefore, that the optimum position of injection can be determined only experimentally by varying the position of injection rearwardly toward the exit and determining the point at which side force passes through its maximum value. Such a position is shown by shock front 20a which is entirely within the nozzle with its rearward terminus intersecting the rear edge of the nozzle and with injection occurring at point 14a.

Tests of angular upstream injection have revealed that while some side force is effected, it is of lesser magnitude than when the liquid is injected perpendicular to the nozzle axis.

The increase in the main thrust of the nozzle appears to increase approximately linearly with the rate of mass flow of the injected liquid which indicates that such increase is due mainly to addition of mass to the main stream.

Various liquids may be employed, however, tests reveal that variations in effectiveness may be expected. For example, water is approximately one-half as effective as N 0 The Freons (e.g. Freon 12) appear to be particularly effective.

It has also been observed that the injection of liquid into the exhaust stream produces a greater side force than is produced by injecting the same liquid through a separate nozzle in an outward direction from the vehicle, a technique formerly employed. Thus, as compared to this former technique, injection of liquid into the nozzle not only produces greater side force but its mass adds to the main thrust of the nozzle.

FIGS. 4 and 5 illustrate a modification employing the invention so far described with additional liquid discharge orifices 14a, 14a and 14b, 14b, disposed in diametrically disposed vanes 22 which produce forces for correcting roll of the missile about its axis when liquid is simultaneously discharged from either pair of orifices 14a, 14a or 14b, 14b, in opposite directions in a plane transverse to the axis. Couples are produced which will move the missile angularly about its axis. As will be understood, the flow of liquid may be controlled by any suitable guidance control which senses roll of the missile about its axis. The phenomenon of the operation is not fully understood a a but it would appear that this modification produces pairs of shock waves which spoil the cross-sectional symmetry of exhaust flow about the nozzle axis, rather than in planes containing the missile axis as previously described in connection with orifices 14. As will also be apparent, orifices 14 are also employed for directional control of the missile along its direction of travel.

Obviously many modifications and variations of the present invention are possible in the light of the above teachings. It is therefore to be understood that within the scope of the appended claims the invention may be practiced otherwise than as specifically described.

What is claimed is:

1. In a rocket motor having a fixed convergent-divergent exhaust nozzle exhausting gas at supersonic velocity and normally producing thrust along the axis of the nozzle, and means for selectively diverting said direction of thrust angularly to said axis in one of a plurality of angularly disposed planes containing said axis, whereby the direction of thrust may be varied as desired and efiect movement of said nozzle in any direction about its axis, said means comprising a plurality of angularly spaced orifices disposed in the divergent portion of said nozzle and in a plane substantially perpendicular to said axis, a source of noncombustible liquid which is unoxidizable by said gas, and means for selectively delivering the liquid to at least one of said orifices for delivery into the exhaust gas in liquid state in suflrcient quantity to produce a planar shock front, having a forward terminus forwardly adjacent the point of delivery and extending angularly rearwardly with a rearward terminus just within the confines of the nozzle.

2. Apparatus in accordance with claim 1 wherein said orifices are disposed at positions along the divergent portion of said nozzle such that the rearward terminus of said shock front substantially intersects a rear edge of said nozzle.

3. Apparatus in accordance with claim 1 including a pair of diametrically disposed radially extending vanes disposed within said divergent portion, and a pair of discharge orifices in each vane adapted to selectively discharge liquid in opposite directions in a plane perpendicular to the axis of the nozzle, and means for simultaneously delivering liquid to one orifice of each pair for delivery in opposite directions to produce other shock fronts for selectively controlling the direction of roll of said nozzle about its axis in either of opposite directions of rotation.

4. A method of controlling the direction of thrust of a convergent-divergent nozzle through which gas is adapted to exhaust at supersonic velocity and normally produce thrust along its axis, comprising, injecting into the exhaust gas a non-combustible liquid which is unoxidizable by said gas, and while in liquid state, in a quantity sufficient for producing a planar shock front having a forward terminus forwardly adjacent the point of delivery and extending angularly rearwardly with a rearward terminus just within the confines of the nozzle, to thereby divert the direction of thrust angularly away from said axis.

5. A method in accordance with claim 4 including the further step of injecting other liquid into said exhaust in two opposite directions spaced from said axis and in a plane perpendicular to said axis to produce additional shock fronts effecting a force couple for rotating said nozzle about its axis.

6. A method in accordance with claim 5 wherein the liquid injected to produce said planar shock front is injected at a point such that the rearward terminus of said planar shock front substantially intersects the rear edge of the nozzle.

References Cited by the Examiner UNITED STATES PATENTS 2,952,123 9/1960 Rich 60-3554 3,066,485 12/1962 Bertin et al. 6035.54

FOREIGN PATENTS 748,983 5/1956 Great Britain.

MARK NEWMAN, Primary Examiner.

ABRAM BLUM, Examiner, 

1. IN A ROCKET MOTOR HAVING A FIXED CONVERGENT-DIVERGENT EXHAUST NOZZLE EXHAUSTING GAS AT SUPERSONIC VELOCITY AND NORMALLY PRODUCING THRUST ALONG THE AXIS OF THE NOZZLE, AND MEANS FOR SELECTIVELY DIVERTING SAID DIRECTION OF THRUST ANGULARLY TO SAID AXIS IN ONE OF A PLURALITY OF ANGULARLY DISPOSED PLANES CONTAINING SAID AXIS, WHEREBY THE DIRECTION OF THRUST MAY BE VARIED AS DESIRED AND EFFECT MOVEMENT OF SAID NOZZLE IN ANY DIRECTION ABOUT ITS AXIS, SAID MEANS COMPRISING A PLURALITY OF ANGULARLY SPACED ORIFICES DISPOSED IN THE DIVERGENT PORITON OF SAID NOZZLE AND IN A PLANE SUBSTANTIALLY PERPENCICULAR TO SAID AXIS, A SOURCE OF NONCOMBUSTIBLE LIQUID WHICH IS UNOXIDIZABLE BY SAID GAS, AND MEANS FOR SELECTIVELY DELIVERING THE LIQUID TO AT LEAST ONE OF SAID ORIFICES FOR DELIVERY INTO THE EXHAUST GAS IN LIQUID STATE IN SUFFICIENT QUANTITY TO PRODUCE A PLANAR SHOCK FRONT, HAVING A FORWARD TERMINUS FORWARDLY ADJACENT THE POINT OF DELIVERY AND EXTENDING ANGULARLY REARWARDLY WITH A REARWARD TERMINUS JUST WITHIN THE CONFINES OF THE NOZZLE. 